1. Field
Disclosed are nozzle configurations for rocket vehicles where the nozzles are not located at the optimal thrust axis of the vehicle. Two examples include nozzles located on the forward end of the vehicle (also called tractor nozzles) and attitude control nozzles located on the periphery of the vehicle. More particularly, the disclosed nozzle shapes enhance the axial thrusts and/or maneuver torques on the vehicle.
2. Description of the Related Art
In conventional rocket nozzle design, the flow and pressure in the nozzle is evaluated as a simple function of the local area to the throat area. FIG. 1 illustrates an axisymmetric nozzle 10 as known from the prior art. The average pressure at each station in the nozzle 10 can be found by inverting the following equation relating the area ratio A/At, pressure ratio Pc/P and specific heat ratio gamma, [7] γ.
                              A          At                =                                            (                                                P                  c                                /                P                            )                                                      γ                +                1                                            2                ⁢                γ                                                                                        [                                                      γ                    +                    1                                    2                                ]                                                              γ                  +                  1                                                  2                  ⁢                                      (                                          γ                      -                      1                                        )                                                                        ⁢                                                            2                                      γ                    -                    1                                                  ⁡                                  [                                                                                    (                                                                              P                            c                                                    /                          P                                                )                                                                                              γ                          -                          1                                                γ                                                              -                    1                                    ]                                                                                        (        1        )            A is local flow area measured in square inches;At is the sonic throat area measured in square inches;Pc is the chamber total pressure measured in pounds per square inch;P is local flow pressure measured in pounds per square inch; andγ is specific heat ratio (unit-less).
Even this calculation has its limits in that the pressure, p, is the average at a station and the wall pressure may be more or less than the average p, a function of Pc & (A/At). This technique has been successfully employed to design axisymmetric nozzles following simple rules. Gas flows into the nozzle 10 at subsonic speed and through a converging portion 12 that terminates at throat 14. The gas then flows at supersonic speed through diverging portion 15 and through exit 16. The supersonic flow in the nozzle 10 responds to changes in the nozzle wall contour 17 through expansion or shock waves. These waves will travel from their origin across the nozzle 10 to the opposite side where they reflect and cross back and forth as the flow accelerates from the throat 14 to the exit 16. These waves are called characteristics or Mach waves and in the days before Computational Fluid Dynamics (CFD), were used to design and analyze the flow in a nozzle 10.
FIG. 2 shows how these Mach waves 18 influence the flow and pressure in the diverging portion 15 of a simplified two dimensional nozzle with a supersonic starting Mach number. This nozzle has a single wall angle change, β, from an axial throat 14 to a 15° expansion. The flow in the nozzle can be broken into several different zones of uniform but strikingly different properties. Zone 0 is the supersonic flow at the throat 14 of the nozzle. An exemplary Mach number is 1.5 and the pressure is 29% of the total pressure of the approaching flow. Zone 1 is downstream of the 15° outward turn. The exemplary Mach number has increased to 1.944 and the pressure has dropped to 14.5%. In Zone 1, the flow properties are uniform and the flow direction is parallel with the nozzle surface. A number of waves can be seen originating at the transition corner 22 and fanning out. This is the well known Prandtl-Meyer Expansion. These Mach waves 18 expand from the origin due to the increasing Mach number as the flow turns and accelerates around the transition corner 22. These waves cross the nozzle centerline 24 generating a Zone 2. In Zone 2, the crossing of the waves from opposite sides of the nozzle cause more turning and acceleration of the flow. The exemplary flow in Zone 2 is now moving parallel to the axis 24 at Mach 2.41 and 6.5% of the total pressure. As the waves continue to crisscross the nozzle, no more Zones of uniform properties are formed in this example and each point in the flowfield has varying Mach number, pressure and flow direction all resulting from the interactions with these waves. If the nozzle is cut off at a fixed length, the average properties at the exit 16 will determine the resulting thrust. Since the flow in not perfectly uniform and axial, a nozzle efficiency term is usually applied for delivered nozzle performance.
In a more general case, the wall contour 17 is not uniform as in this example and the pressure acting on the wall is affected directly by changes in wall angle and waves from the opposite side of the nozzle. Since the Mach waves move at the local speed of sound in the gas, they move away from the wall and are swept downstream by the supersonic flow. This means that wall angle changes affect the local pressure and the pressure on the opposite wall well downstream of the origin of that change. If, for example, the lower wall 26 was terminated at point A, the flow would expand out and downward causing another expansion fan to travel from A to B. Since B is beyond the exit of the nozzle the removal of this lower wall section will not change the pressures on the upper wall 28.
In most rocket propelled vehicles, one or more rocket nozzles are located at the aft end and aligned close to the vehicle axis to convert most of the thrust in flight direction. Some applications exist that preclude this aft location for nozzles. Exemplary tractor nozzle applications are the TOW missile and the escape motor for the Orion, Crew Exploration Vehicle. In another application, multiple rocket nozzles are arrayed around the vehicle for maneuvering. For both of these applications, using conventional nozzles can have a detrimental effect on vehicle performance.
Referring to FIG. 3, the launch abort system 30 of an Orion Crew Exploration Vehicle 32, or similar rocket, has a payload 34 that includes a crew module 36 and a service module 38 located forward of booster rocket 40.
In the event that it becomes necessary to abort the mission prior to separation of the booster rocket 40 from the payload 34, abort motor 42 is ignited generating propellant gases that are expelled through nozzles 44. Nozzles 44 are angularly disposed in an aftward direction to the intended direction of flight of the vehicle and generate a thrust effective to separate the launch abort system and crew module 30 from the remainder of the vehicle 32. This propulsion system has a number of limitations. Hot propellant gases expelled by the nozzles 44 impact the abort motor 42 and crew module 36. These components must be designed to withstand the high temperatures that may be generated by impingement of the propellant gases. One approach, that was utilized on the Apollo program, is to cant the nozzles 44 at an extreme angle to expel the hot propellant gases sufficiently outward from the rocket components to avoid the most severe temperature increases. However, deviation of the nozzle direction from directly aftward causes a loss of thrust requiring abort motor 42 to be charged with additional propellant.
These limitations are illustrated in more detail in FIGS. 4 and 5. Referring to FIG. 4, when the nozzle 44 is canted at an angle, α, relative to direction of flight 26, the cant angle on the multiple nozzles causes a loss in thrust that is roughly equal to (1−cos α) which at 30 degrees of cant represents a thrust loss of 13%.
Nozzles 44 project outward from exterior walls 48 of the abort motor 42 that arrayed around motor centerline 47. The abort motor 42 includes a propellant 49 that when ignited generates propellant gases. The propellant gases pass through nozzle throat 14 and are accelerated through divergent portion 15 of the nozzle. If the cant angle, α, of the nozzle is greater than the divergence angle, β, the wall 50 of the nozzle 44 generates a negative thrust due to the local nozzle pressures acting in a rearward direction. The opposite effect is seen on wall 53. Here the wall pressures 55 have a larger forward projected area and contribute more to the net thrust of the nozzle 44. However, with the conventional nozzle 44, the thrust is significantly less than an axially directed nozzle (due to the cosine loss from the previous paragraph).
Referring now to FIG. 5, the abort motor 42 must function from sea level static to very high altitudes. The most important condition is likely to be at the maximum dynamic pressure (max Q) point in the assent trajectory. This is the point of highest drag and stress on the vehicle and is the most challenging for the design of the escape system. Under these conditions, the approaching supersonic flow 52 causes a bow shock wave 54 and elevated pressures that deflect the rocket plume 56 back towards the motor exterior walls 48 and crew module 36. With reference to FIG. 6, this results in impingement of underexpanded rocket plume 56 on the exterior wall 48 at attachment point 58. The plume impingement separates the boundary layer 60 which on reattachment, results in a hot spot that may degrade the exterior wall 48 structural capability.
Referring back to FIG. 5, making the situation worse is the fact that the attachment point 58′, 58″, 58′″ shifts with flight speed. A second problem is that despite a large cant angle, α, the exhaust plume 56′ still interacts with the crew module 36. Increasing the distance between the nozzles 44 and crew module 36 will reduce the heating caused by the plume 56′ but will not avoid the flow reaching the forward surface of the crew module.
The crew module 36 typically includes roll control nozzles 62 that control module roll while in orbit. As shown in FIG. 7, the roll control nozzle 62 has an exit portion that is flush with the perimeter 64 of the crew module 36 or other space vehicle. The function of this rocket thruster is to control the orientation of the vehicle by providing a torque about its center of mass (CG). A conventional roll control nozzle 62 is installed with one exit side 65 flush with the vehicle perimeter 64, the nozzle is then scarfed so that the opposing exit side 67 is also flush with the vehicle perimeter (no portion of the nozzle protrudes from the vehicle). The expanding gases in the nozzle apply a force F1 along the nozzle axis and also F2 normal to the nozzle axis. The intended roll controlling torque is F1*R1 while the scafing creates a counter torque (F2*R2) on the unbalanced surface A2. The net torquer, applied to the vehicle is then:τ=F1*R1−F2*R2  (2)Compensation for the counter torque requires more propellant to be consumed for a given maneuver than desired.
There remains a need for improved rocket nozzles, such as for a launch abort system, or the capsule attitude control system, that overcomes this loss of vehicle performance.